In a combustion system for a gas turbine, fuel and compressed air are mixed together and ignited to produce hot combustion gases that drive a turbine and produce thrust or drive a shaft coupled to a generator for producing electricity. In an effort to reduce pollution levels, government agencies have introduced new regulations requiring gas turbine engines to reduce emitted levels of emissions, including carbon monoxide (CO) and oxides of nitrogen (NOx). A common type of combustion, employed to comply with these new emissions requirements, is premix combustion, where fuel and compressed air are mixed together prior to ignition to form as homogeneous a mixture as possible and burning this mixture to produce lower emissions. While premixing fuel and compressed air prior to combustion has its advantages in terms of emissions, it also has certain disadvantages such as combustion instabilities and more specifically combustion dynamics.
In order to achieve the lowest possible emissions through premixed combustion, without the use of a catalyst, it is necessary to provide a fuel-lean mixture to the combustor. Generally, in combustors using fuel-lean mixtures, as the fuel-air mixture becomes more fuel rich (i.e. higher fuel-air ratio), the flame and combustion process becomes more stable. (Of course, if the fuel-air mixture becomes too fuel rich, the flame and combustion process becomes more unstable, and can lead to rich fuel blowout.) Therefore, fuel-lean mixtures tend to be more unstable given the lesser fuel content for a given amount of air. As a result, when fuel-lean mixtures are burned they tend to produce greater pressure fluctuations due to the unstable flame. A factor contributing to the unstable flame is the fuel-air ratio or more specifically, the amount of air mixing with a known amount of fuel. The amount of air entering into a combustion chamber can vary depending on how the air is directed towards the combustion chamber inlet. If the airflow is not uniform and not relatively free from swirl, the amount of air entering the combustor will fluctuate, thereby altering the fuel-air ratio, and adversely affecting combustion stability and NOx emissions.
An example of a gas turbine combustor of the prior art that employs premix combustion is shown in cross section in FIG. 1. A gas turbine combustor 10 comprises fuel injection system 11, combustion liner 12, transition duct 13, first outer sleeve 14, and second outer sleeve 15. For the combustor shown in FIG. 1, air used for combustion, represented by arrows, enters into generally annular passage 16 through a plurality of holes in first outer sleeve 14 and second outer sleeve 15. In this prior art system, the air enters at different axial locations and at different angles, including generally perpendicular to the walls of combustion liner 12 and transition duct 13. As a result, the air flow in generally annular passage 16 has some swirl, or tangential velocity component. It is this swirl that causes a non-uniform air flow distribution to combustion liner 12, and hence creates combustion stability problems by causing the fuel-air ratio in the combustor to fluctuate. In order to try and non-mechanically reduce the swirl effects, a greater pressure drop was taken across generally annular passage 16 through the sizing of passage 16 and sizing of plurality of holes in first outer sleeve 14 and second outer sleeve 15. The additional pressure drop taken across the combustor of the prior art results in overall loss in the efficiency of the gas turbine. As those skilled in the art will readily appreciate, higher pressure drop across the combustion system results in lower gas turbine cycle efficiency, so designers of gas turbine combustion systems seek to minimize this pressure drop.
Therefore, it is desired to provide a combustion system for a gas turbine wherein the geometry of the combustor provides a means for significantly reducing the tangential velocity, or swirl, for air directed to a combustion inlet so as to reduce combustion stability problems and NOx, and to reduce the overall pressure drop required across the combustor. Reducing the combustor pressure drop, will in turn improve the gas turbine efficiency, increase power output, and lower fuel operating cost.